The invention relates to supersonic diffusers. The invention relates more particularly to external-compression supersonic diffusers such as are used as inlets in air-breathing propulsion systems.
In a variety of types of aircraft, air-breathing propulsion systems such as turbojet or turbofan engines are used for propelling the aircraft at supersonic velocities. Existing commercially available gas turbine engines used for aircraft propulsion are virtually invariably designed to work in a regime in which subsonic flow, typically on the order of Mach 0.3 to 0.6, exists at the upstream face of the engine. Thus, a supersonic diffuser or inlet is necessary to decelerate the captured supersonic air stream to a subsonic speed for ingestion by the engine. The process of deceleration is technically known as diffusion or compression, since the excess kinetic energy of the air stream is converted into a static pressure increase. To maximize the overall propulsive efficiency of the engine/inlet system, the inlet must perform its diffusion function efficiently. The efficiency of the diffusion process is a function of how much total pressure is lost in the air stream between the entrance side of the inlet and the discharge side. The total-pressure recovery of an inlet is defined by a ratio of total pressure at the discharge to total pressure at the entrance. A primary objective of inlet design is to maximize total pressure recovery. External drag on the inlet also affects the overall efficiency of the system, and thus it is desirable to minimize such drag. Additionally, a further objective of the inlet design process is to maximize flow stability so as to avoid violent flow oscillations, primarily unstart, that can occur with some types of supersonic inlets.
A supersonic inlet generally includes a forward portion comprising a converging supersonic diffuser, and an aft portion comprising a diverging subsonic diffuser. Most supersonic inlets are either two-dimensional or xe2x80x9c2Dxe2x80x9d having a rectangular-shaped flow area, or axisymmetric having a circular flow area. A throat of the inlet occurs at the juncture between the supersonic diffuser and the subsonic diffuser where the flow area reaches a minimum. Supersonic inlets are generally classified into three types: internal compression, mixed compression, and external compression. Internal-compression inlets are designed to accomplish both supersonic and subsonic compression within the interior of the inlet duct, and thus the shock structure of the supersonic compression field must be xe2x80x9cswallowedxe2x80x9d into the inlet duct in order for the inlet to work as designed. The problem of xe2x80x9cunstartxe2x80x9d occurs in an internal-compression inlet when a flow disturbance causes the terminal shock to be expelled out the forward end of the inlet duct. The result is a drastic loss in efficiency and a large increase in inlet drag. Unstart thus represents a significant problem.
Mixed-compression inlets are those in which part of the supersonic compression is accomplished forward of the inlet duct aperture by forcing the approaching air stream to turn prior to being ingested into the duct. Supersonic compression continues internally in the forward part of the duct, followed by subsonic compression. These types of inlets can still suffer from flow instability problems such as unstart, since the terminal shock must still be swallowed into the duct as with internal-compression inlets.
External-compression inlets accomplish all supersonic compression externally such that the flow in the inlet duct is all subsonic. External-compression inlets are less susceptible to unstart-type instabilities because the terminal shock tends to remain stable in its position at the entrance to the inlet duct, which represents the throat of the inlet. However, external-compression inlets are typically disfavored for flight above about Mach 2.0 because they tend to have high cowl drag as a consequence of the large amount of flow turning that must be accomplished forward of the inlet duct. This large flow turning leads to high cowl angles and long cowl lengths in the cross-stream direction, and thus high drag.
It would be desirable to provide a supersonic inlet having good flow stability such as that typical of conventional external-compression inlets, and at the same time having high total-pressure recovery and low external drag.
The above needs are met and other advantages are achieved by the present invention, which provides an external-compression inlet having a unique three-dimensional external compression surface that can provide high total pressure recovery and low cowl drag while maintaining the good flow stability that is characteristic of external-compression inlets. To these ends, a supersonic external-compression inlet in accordance with a preferred embodiment of the invention comprises a generally scoop-shaped supersonic compression section for diffusing a supersonic free stream flow. The supersonic compression section includes a main wall having a leading edge and a throat portion downstream of the leading edge, and side portions joined to opposite side edges of the main wall so as to form a generally scoop-shaped structure. The side portions advantageously extend into the supersonic flow stream far enough to encompass the initial oblique shock wave that is attached to the leading edge of the main wall. The main wall has an inner surface formed generally as an angular sector of a surface of revolution, the inner surface of the main wall coacting with inner surfaces of the side portions to define a three-dimensional external-compression surface. The supersonic external-compression inlet also includes a subsonic diffuser section arranged to receive flow from the supersonic compression section and to diffuse the flow to a subsonic condition. The subsonic diffuser section is formed by a cowl shaped as a closed duct, the cowl having a leading-edge cowl lip spaced in a cross-stream direction from the throat portion of the main wall such that a throat of the supersonic inlet is defined proximate the cowl lip between the cowl and the throat portion. The three-dimensional external-compression surface of the inlet enables the external flow turning to be reduced relative to a conventional 2D or axisymmetric supersonic inlet, and accordingly the external cowl angle can be reduced. The unique scoop-shaped supersonic diffuser section also enables the cowl length in the cross-stream direction to be reduced relative to a diffuser of conventional type, because the three-dimensional compression surface does not completely surround the flow stream. Drag on the scoop-shaped diffuser section can thus be reduced relative to a conventional supersonic diffuser.
The invention also encompasses a variable-geometry inlet of simple construction. In accordance with a preferred embodiment of the invention, the main wall of the supersonic compression section includes a movable external ramp that pivots about its forward edge. The aft portion of the external ramp defines the throat portion of the main wall. Thus, the size of the throat can be varied by pivoting the ramp so as to vary the distance between the throat portion and the cowl lip. The cowl preferably also includes a movable internal ramp located aft of the throat. The internal ramp is pivotable about its aft edge and has a forward edge that is proximate the aft edge of the external ramp. The internal ramp is pivoted in concert with the external ramp so as to maintain a smooth flow transition therebetween. Advantageously, the external and internal ramps are formed by simple hinged plates. If desired, the external and internal ramps can be spaced apart slightly in the flow direction where they meet so as to create a slot that can be used for boundary layer bleed.
Various configurations of three-dimensional external-compression surfaces can be used in accordance with the present invention. Preferably, however, all such configurations should be shaped in accordance with a design method of the invention, in which the external-compression surface comprises a surface that is fit through streamlines that originate at the perimeter of an upstream capture area for the inlet. First, an axisymmetric design is performed to determine an axisymmetric compression surface having a compression field providing a compression field that yields good total pressure recovery. Next, the shape of the inlet capture area is prescribed such that the portion of the axisymmetric compression field that yields good pressure recovery is captured by the capture area. A portion of the capture area""s perimeter is defined by an angular sector of the axisymmetric compression surface at the leading edge of the supersonic compression section. Streamlines from the axisymmetric flow solution are traced from a plurality of points located about the perimeter of the capture area. A surface fit through these streamlines defines the three-dimensional external-compression surface for the inlet.
In one embodiment of the invention, the inner surface of the main wall has a circular-arc shape in cross-section normal to the free-stream direction and subtends a constant circular-arc angle from the leading edge to the throat. Advantageously, the side portions of the supersonic compression section comprise two substantially planar side walls respectively joined to opposite side edges of the main wall and extending generally radially with respect to the circular-arc inner surface thereof, the side walls extending from the leading edge of the main wall to the cowl lip of the subsonic diffuser section and being joined to the cowl lip. Preferably, the supersonic compression section at a discharge end thereof defines a flow area configured as a sector of an annulus, the subsonic diffuser section at an inlet end thereof defines a flow area configured to substantially match that of the discharge end of the supersonic compression section, and the discharge end of the subsonic diffuser section defines a substantially circular flow area. The subsonic diffuser section desirably provides a smooth transition from the annulus sector flow area at its entrance to the circular flow area at its exit.
In accordance with a particularly preferred embodiment of the invention, the inner surface of the main wall is contoured in the flow direction so as to create an initial weak oblique shock wave at the leading edge of the diffuser followed by isentropic compression to a Mach number of about 1.3 at the throat. A series of Mach lines (shock waves of virtually no strength such that substantially no pressure loss occurs across them) radiate from the inner surface. The inner surface is designed such that at a predetermined flow condition the initial weak shock wave and the Mach lines all intersect at a common focal point spaced in the cross-stream direction from the throat portion of the main wall. The cowl lip is located substantially at the common focal point. This design enables the spillage drag of the inlet to be minimized by ensuring that the subsonic diffuser captures all or nearly all of the externally compressed flow.
The invention also encompasses a supersonic external-compression inlet integrated into an aircraft wing in an advantageous manner. The shape of the local wing surface is modified to fit part or all of the supersonic diffuser contour generated by the design procedure of the invention. The supersonic diffuser wall thus serves both as the diffuser surface and as the local wing surface. This leads to reduction in wetted surface area relative to an alternative design in which the wing surface and diffuser surface are separate members. The reduction in wetted area in turn provides reduced skin friction drag for the inlet/wing system.
The invention thus provides a unique supersonic diffuser, and method for designing such a diffuser, enabling a compression field substantially duplicating that of an axisymmetric compression surface to be produced with a scoop-shaped supersonic compression section that does not completely surround the flow stream. Accordingly, cowl drag can be reduced in view of the reduced cowl wetted area relative to an axisymmetric cowl. The short cowl length of the supersonic compression section also leads to reduced weight for the diffuser. Total pressure recovery of the diffuser is predicted by CFD modeling to be equal to the highest levels obtained in wind tunnel tests of mixed-compression inlets designed for the same flight Mach number. The diffuser is also predicted to provide the good flow stability that is typical of external-compression inlets.